CA2904200A1 - Dual-wall impingement, convection, effusion combustor tile - Google Patents

Dual-wall impingement, convection, effusion combustor tile Download PDF

Info

Publication number
CA2904200A1
CA2904200A1 CA2904200A CA2904200A CA2904200A1 CA 2904200 A1 CA2904200 A1 CA 2904200A1 CA 2904200 A CA2904200 A CA 2904200A CA 2904200 A CA2904200 A CA 2904200A CA 2904200 A1 CA2904200 A1 CA 2904200A1
Authority
CA
Canada
Prior art keywords
tile
gas turbine
wall
combustor
liner
Prior art date
Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
Abandoned
Application number
CA2904200A
Other languages
French (fr)
Inventor
Jack D. Petty
Mohan RAZDAN
Michael S. Bell
Michel S. Smallwood
Tab M. Heffernan
Current Assignee (The listed assignees may be inaccurate. Google has not performed a legal analysis and makes no representation or warranty as to the accuracy of the list.)
Rolls Royce Corp
Rolls Royce North American Technologies Inc
Original Assignee
Rolls Royce Corp
Rolls Royce North American Technologies Inc
Priority date (The priority date is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the date listed.)
Filing date
Publication date
Application filed by Rolls Royce Corp, Rolls Royce North American Technologies Inc filed Critical Rolls Royce Corp
Publication of CA2904200A1 publication Critical patent/CA2904200A1/en
Abandoned legal-status Critical Current

Links

Classifications

    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
    • F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
    • F23R3/00—Continuous combustion chambers using liquid or gaseous fuel
    • F23R3/002—Wall structures
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
    • F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
    • F23R3/00—Continuous combustion chambers using liquid or gaseous fuel
    • F23R3/007—Continuous combustion chambers using liquid or gaseous fuel constructed mainly of ceramic components
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
    • F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
    • F23R3/00—Continuous combustion chambers using liquid or gaseous fuel
    • F23R3/02—Continuous combustion chambers using liquid or gaseous fuel characterised by the air-flow or gas-flow configuration
    • F23R3/04—Air inlet arrangements
    • F23R3/06—Arrangement of apertures along the flame tube
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
    • F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
    • F23R3/00—Continuous combustion chambers using liquid or gaseous fuel
    • F23R3/42—Continuous combustion chambers using liquid or gaseous fuel characterised by the arrangement or form of the flame tubes or combustion chambers
    • F23R3/60—Support structures; Attaching or mounting means
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2230/00—Manufacture
    • F05D2230/20—Manufacture essentially without removing material
    • F05D2230/21—Manufacture essentially without removing material by casting
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2230/00—Manufacture
    • F05D2230/20—Manufacture essentially without removing material
    • F05D2230/22—Manufacture essentially without removing material by sintering
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2260/00—Function
    • F05D2260/20—Heat transfer, e.g. cooling
    • F05D2260/201—Heat transfer, e.g. cooling by impingement of a fluid
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2260/00—Function
    • F05D2260/20—Heat transfer, e.g. cooling
    • F05D2260/202—Heat transfer, e.g. cooling by film cooling
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
    • F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
    • F23R2900/00—Special features of, or arrangements for continuous combustion chambers; Combustion processes therefor
    • F23R2900/03041—Effusion cooled combustion chamber walls or domes
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
    • F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
    • F23R2900/00—Special features of, or arrangements for continuous combustion chambers; Combustion processes therefor
    • F23R2900/03042—Film cooled combustion chamber walls or domes
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
    • F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
    • F23R2900/00—Special features of, or arrangements for continuous combustion chambers; Combustion processes therefor
    • F23R2900/03043—Convection cooled combustion chamber walls with means for guiding the cooling air flow
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
    • F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
    • F23R2900/00—Special features of, or arrangements for continuous combustion chambers; Combustion processes therefor
    • F23R2900/03044—Impingement cooled combustion chamber walls or subassemblies
    • Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
    • Y02—TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
    • Y02T—CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
    • Y02T50/00—Aeronautics or air transport
    • Y02T50/60—Efficient propulsion technologies, e.g. for aircraft

Landscapes

  • Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Ceramic Engineering (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)

Abstract

A gas turbine engine includes a combustor having a dual-wall impingement convention effusion combustor tile assembly. The dual-wall tile assembly provides a cooling air flow channel and attachments for securing the tile to the cold skin liner of the combustor. Cooling is more efficient in part due to the dual wall construction and in part due to reduced parasitic leakage, and the design is less sensitive to attachment features which operate at lower temperatures.

Description

DUAL-WALL IMPINGEMENT, CONVECTION, EFFUSION COMBUSTOR TILE
CROSS-REFERENCE TO RELATED APPLICATIONS
[0001] This application claims priority to U.S. Provisional Patent Application No.
61/773,082 filed March 5, 2013, the contents of which are hereby incorporated in their entirety.
FIELD OF TECHNOLOGY
[0002] A gas turbine engine uses a combustor and a combustor liner, and more particularly, a liner having wall elements to form a dual wall cooling system.
BACKGROUND
[0003] Gas turbine engines are used extensively in high performance aircraft and they employ fans, compressors, combustors and turbines and during operation they generate energies and air flows that impact the performance of the engine's systems. A
gas turbine may employ one or more combustors that serve as the fuel preparation and ignition chambers for generating the temperature rise which is required to drive the turbine blades. Typical combustors may use inner and outer liners that define an annular combustion chamber in which the fuel and air mixtures are combusted. The inner and outer liners are radially offset from the combustor casings such that inner and outer passage ways are defined between the respective inner and outer liners and casings.
[0004] In order to improve the thrust and fuel consumption of gas turbine engines, i.e., the thermal efficiency, it is necessary to use high compressor exit pressures and combustion exit temperatures. Higher compressor pressures also give rise to higher compressor exit temperatures supplied to the combustion chamber, which results in a combustor chamber experiencing much higher temperatures than are present in most conventional combustor designs.
[0005] A need exists to provide effective cooling of the combustion chamber walls.
Various cooling methods have been proposed including the provision of a doubled walled combustion chamber whereby cooling air is directed into a gap between spaced outer and inner walls, thus cooling the inner wall. This air is then exhausted into the combustion chamber through apertures in the inner wall. The inner wall may be comprised of a number of heat resistant tiles.
[0006] Combustion chamber walls which comprise two or more layers are advantageous in that they only require a relatively small flow of air to achieve adequate wall cooling.
However, hot spots may form in certain areas of the combustion chamber wall.
This problem is heightened as temperatures within the combustion chamber which can exceed 3,500 degrees F. Such harsh environmental conditions may prematurely reduce the life of the liner of the combustor. In addition, loss of tile attachment and subsequent component distress remains an engineering challenge in current combustor technology.
BRIEF DESCRIPTION OF THE DRAWINGS
[0007] While the claims are not limited to a specific illustration, an appreciation of the various aspects is best gained through a discussion of various examples thereof Referring now to the drawings, exemplary illustrations are shown in detail. Although the drawings represent the illustrations, the drawings are not necessarily to scale and certain features may be exaggerated to better illustrate and explain an innovative aspect of an example. Further, the exemplary illustrations described herein are not intended to be exhaustive or otherwise limiting or restricted to the precise form and configuration shown in the drawings and disclosed in the following detailed description. Exemplary illustrations are described in detail by referring to the drawings as follows:
[0008] FIG. 1 illustrates a schematic diagram of a gas turbine engine employing an improved combustor assembly;
[0009] FIG. 2 illustrates a side sectional view of a gas turbine engine with an improved tiled combustor assembly;
[0010] FIG. 3 illustrates a partial perspective sectional view of a gas turbine engine with a tiled combustor assembly;
[0011] FIG. 4 illustrates a partial sectional view of a combustor assembly showing the installation of a DICE tile;
[0012] FIG. 5 illustrates a perspective view a Dual-Wall Impingement, Convection, Effusion (DICE) combustor tile, showing the hot side and a cold side;
[0013] FIG. 6 illustrates an enlarged perspective view of the cold side of the DICE tile, showing a cut away depicting the pedestals and air channels; and
[0014] FIG. 7 illustrates an enlarged side cross-sectional view of a DICE tile assembly for use in a combustor of a gas turbine engine.
DETAILED DESCRIPTION
[0015] A gas turbine engine combustor tile design includes an exemplary high temperature capable dual wall combustor tile attached to a lower temperature capable cold skin of a combustor liner. The wall cooling is accomplished by feeding air through holes in the cold skin. The air impinges on the back side of the hot tile and then flows out ejection slots or holes into the hot flow path. The gap formed between the cold skin and the tile hot side surface forms a cooling channel which may be enhanced by the presence of turbulators or pin fins. This interface gap is maintained by pulling the hot tile into the cold skin via attachment features such as studs. Standoffs on the back side of the tile land against the cold skin and react against the fastener preload in order to maintain position of the tiles during engine operation.
[0016] The exemplary tile assembly 42 is a dual-wall impingement, convection, effusion combustor tile and method of constructing a tile which offers significant benefit over conventional combustor wall cooling systems in terms of temperature capability and cooling flow requirements. The embodiment disclosed herein blends the technology of a tiled combustor liner with an integral dual wall cooling system to form a novel tile assembly.
[0017] Figure 1 illustrates a gas turbine engine 10, which includes a fan 12, a low pressure compressor and a high pressure compressor, 14 and 16, a combustor 18, and a high pressure turbine, intermediate pressure, and low pressure turbine, 20 thru 22, respectively. The high pressure compressor 16 is connected to a first rotor shaft 24, the low pressure compressor 14 is connected to a second rotor shaft 26, and the fan 12 is connected to a third rotor shaft 43.
The shafts extend axially and are parallel to a longitudinal center line axis 28. It will be appreciated that the improvements disclosed herein can be used with gas turbine engines that incorporate a single or two-shaft architecture.
[0018] Ambient air 30 enters the fan 12 and is directed across a fan rotor 32 in an annular duct 34, which in part is circumscribed by fan case 36. The bypass airflow 38 provides engine thrust while the primary gas stream 40 is directed to the compressors 14 and 16, combustor 18, and the turbines 20 thru 22. The gas turbine engine 10 includes an improved combustor 18 having a tile assembly 42, the details of the exemplary design are set forth herein.
[0019] FIG. 2 illustrates a side sectional view of the combustor 18 with a plurality of tile assemblies 42 that are secured to a cold skin or outer surface of a liner 44.
A combustor outer case 46 circumscribes a combustor shell 48 and a fuel nozzle 50 provides pressurized fuel 52 to a combustor chamber 54. The combusted fuel may be ignited by an igniter (not shown) which in turn subjects the chamber 54 to elevated temperatures which can exceed 3,500 degrees F. Such arrangement causes extreme temperatures to impinge upon the hot surface 56 of each tile assembly 42. A fastener 60 secures each tile assembly 42 to the liner 44 of the combustor 18. The tile assembly 42 is serviceable and may be replaced when it is damaged or is otherwise sufficiently depleted in performance quality.
[0020] FIG. 3 illustrates the shell 48 of the combustor 18 having a plurality of tile assemblies 42 spaced apart and secured to the inner surface 58 of the skin 44.
The inner surface 58 is protected by the tile assembly 42 at substantially the entire inner surface 58 of the skin 44. A gap 60 is maintained between the inner surface 58 and the assembly 42. The cooling effectiveness of each dual wall tile assembly 42 does not rely on accurately maintaining the gap 60 between the tile standoff features and the cold skin 44, as is the case for conventional tiles. In addition, the tile attachment feature or fastener 60 will be maintained at a lower temperature as compared to a conventional tile system.
This arrangement results in a robust mechanical attachment that resists creep and loss of preload, both of which translate into improved component reliability/durability and reduced parasitic leakage. Parasitic leakage which bypasses the cooling circuit translates into lower overall cooling effectiveness.
[0021] Reduced combustor wall cooling translates into a competitive advantage in term of combustor pattern factor control, radial temperature profile control, efficiency, and emissions reduction. The integral dual wall metallic combustor tile assembly 42 offers significant advantages over conventional tiles including but not limited to a reduction in wall cooling flow, a cooler tile attachment (improved reliability/durability), reduced tile leakage and the associated penalty in cooling effectiveness due to leakage, and a more robust mechanical design in terms of less sensitivity to cold skin and tile geometric tolerances/operating deflections.
[0022] FIG. 4 illustrates a cut away of the combustor 18 showing one tile assembly 42 shown offset from the cold skin inner surface 58 of a combustor 18. A tile mounting surface 62 on the cold skin inner surface 58 provides a mounting space for receiving each tile assembly 42. The tile assembly 42 is shown offset from the surface 62 for illustrative purposes. The tile mounting surface 62 has substantially the same profile as the profile of the tile assembly 42. The mounting surface 62 has a plurality of apertures 64 for providing cooling air flow. Small and large dilution ports, 66 and 68, provide large airflow passageways through the skin 44. Surface 70 of the tile assembly 42 represents the hot side of the tile which is subjected to extreme heat conditions.
[0023] FIG. 5 illustrates an exploded view showing a first wall or cold side 72 and a second wall or hot side 70 of the tile assembly 42. The walls 70 and 72 may be substantially planner or of high curvature in configuration. The cold side 72 and the hot side 70 are shown split apart for illustration purposes only. The hot side 70 represents a front side of the assembly 42 and the cold side 72 represents a back side of the assembly 42.
The assembly 42 could be constructed from metal or a composite ceramic material.
[0024] The hot side 70 includes cooling exit holes or slots 74, small dilutions holes 76, and large dilution holes 78. The cold side 72 of the tile assembly 42 includes cooling entry holes 73, and co-aligned small dilution ports 76 and large dilution ports 78. . The hot side of the tile assembly 42 also includes a plurality of cooling exit holes 74. A
plurality of threaded studs or fasteners 60 extend from a surface 80 of the first wall 72. A rail or lip 82 protrudes from the surface 80 around the perimeter of the first wall 72 and is rhombus shaped but other shapes are contemplated. The rail 82 may be integral with the surface 80. A
surface of the rail 82 impinges upon the inner surface 58 of the cold skin 44. The rail creates a plenum 92 to feed the cooling holes 74 and operates to create an offset from a surface of the cold skin.
[0025] FIG. 6 illustrates an enlarged perspective view of the first wall or cold side 72 which is the cold side of the assembly 42. A cut away section 84 is depicted in the lower portion of FIG. 6 which illustrates, under the outer surface 80, a plurality of square-shaped pedestals 86 that are offset by air channels 88. The pedestal pattern 90 consisting of the pedestals 86 and air channels 88 shown is exemplary in nature and other geometric configurations are contemplated. The pattern 90 extends underneath substantially the entire surface 80 and provides air flow channels 88 for aiding cooler air distribution about the first and second walls 70 and 72.
[0026] FIG. 7 illustrates a cross-sectional view taken from line 7-7 of FIG.
2, depicting a tile assembly 42 secured to a cold skin 44. The tile assembly 42 may be constructed primarily of a composite ceramic material (CMC), but other configurations could include a metallic two-piece diffusion or braze bonded assembly of cast, wrought, or direct metal laser sintered (a/k/a direct laser deposition or additive manufactured) components, or a single piece cast or direct metal laser sintered tile. The tile's hot surface can either be as manufactured or can have a thermal and/or environmental barrier coating applied. The coating could be ceramic. The cross-section that is shown in FIG. 7 includes a stud 60 extending through the cold skin 44 of the combustor. A nut or other anchor 61 can be provided as well so as to provide a mechanical securing mechanism for attaching each assembly 42 to the skin 44. The cool side of the DICE tile assembly 42 has a rail 82 upwardly impinging upon the underside 58 of the cold skin 44, thus creating a plenum 92. The wall of the cold side 72 is offset from the wall 70 of the hot side by pedestals 86, the distance of which can be modified to enhance air channel 88 capacities and volumes.
Normal, angled, and/or shaped cooling holes 74 may extend from the air channels 88, through the hot side wall 72, and then into the interior 54 of the combustion chamber 18.
[0027] It will be appreciated that the aforementioned method and devices may be modified to have some components and steps removed, or may have additional components and steps added, all of which are deemed to be within the spirit of the present disclosure. Even though the present disclosure has been described in detail with reference to specific embodiments, it will be appreciated that the various modifications and changes can be made to these embodiments without departing from the scope of the present disclosure as set forth in the claims. The specification and the drawings are to be regarded as an illustrative thought instead of merely restrictive thought.

Claims (20)

What is claimed is:
1. A gas turbine engine having a combustor comprising:
a liner having an inner surface;
a plurality of tile assemblies secured to the inner surface of the liner, each tile assembly having a cooling plenum; and at least one securing member for securing each tile assembly to the liner.
2. The gas turbine as claimed in claim 1, wherein the cooling plenum feeding the tile assembly includes a first wall and a rail, the rail engages a surface of the liner to form a space for cooling air to flow.
3. The gas turbine as claimed in claim 1, wherein the tile assembly includes a first wall and a second wall that are separated by at least one pedestal member.
4. The gas turbine as claimed in claim 1, wherein the tile assembly includes individual members that are formed of a ceramic composite material.
5. The gas turbine as claimed in claim 1, wherein the tile assembly includes a single member formed of a ceramic composite material.
6. The gas turbine as claimed in claim 1, wherein the tile assembly includes individual members that are formed of a metallic material.
7. The gas turbine as claimed in claim 1, wherein the tile assembly includes a first substantially planar or curved member, a second substantially planar or curved member, a spacer for offsetting said members, and apertures passing through each member for providing a cooling flow path.
8. The gas turbine as claimed in claim 7, wherein the securing member is connected to one of the planar or curved members.
9. The gas turbine as claimed in claim 1, wherein each tile assembly includes a plurality of securing members.
10. The gas turbine as claimed in claim 1, wherein each tile assembly is constructed of one or more ceramic matrix members.
11. The gas turbine as claimed in claim 1, wherein each tile assembly includes a first and second wall, each said wall includes a small dilution port and/or a large dilution port.
12. The gas turbine as claimed in claim 1, wherein the inner case has a tile mounting surface, the tile mounting surface includes cooling ports that permit air flow to the tile assembly.
13. A tile for a gas turbine engine combustor comprising:
a first wall element;
a second wall element; and a cooling plenum disposed between the first and second wall elements.
14. The tile for a gas turbine engine as claimed in claim 12, further comprising a mounting stud that is secured at one end to one of the wall elements, the mounting stud is secured at another end to a wall of a combustor liner.
15. The tile for a gas turbine engine as claimed in claim 12, wherein the first wall element includes a plurality of air channels that are separated by pedestals.
16. The tile for a gas turbine engine as claimed in claim 12, wherein the first wall element includes a rail that extends along an outer perimeter of the wall element.
17. The tile for a gas turbine engine as claimed in claim 12, wherein the second wall element includes cooling holes that extend through the second wall element for delivering air to the first wall element.
18. The tile for a gas turbine engine as claimed in claim 12, further comprising a combustor liner, the first wall element is positioned adjacent to the combustor liner.
19. A combustor for a gas turbine engine comprising:
a liner;
a multi-walled tile, the tile haying a first layer, a second layer, a stud, a dilution port, an air plenum disposed between the first and second layer; and a fastening member for securing the tile to the liner.
20. The combustor for a gas turbine engine as claimed in claim 18, wherein the first layer has an outside surface with cooling holes, and air channels in a section underneath the first surface, the air channels are separated by members that route air underneath the outside surface.
CA2904200A 2013-03-05 2013-12-03 Dual-wall impingement, convection, effusion combustor tile Abandoned CA2904200A1 (en)

Applications Claiming Priority (3)

Application Number Priority Date Filing Date Title
US201361773082P 2013-03-05 2013-03-05
US61/773,082 2013-03-05
PCT/US2013/072931 WO2014137428A1 (en) 2013-03-05 2013-12-03 Dual-wall impingement, convection, effusion combustor tile

Publications (1)

Publication Number Publication Date
CA2904200A1 true CA2904200A1 (en) 2014-09-12

Family

ID=50943516

Family Applications (1)

Application Number Title Priority Date Filing Date
CA2904200A Abandoned CA2904200A1 (en) 2013-03-05 2013-12-03 Dual-wall impingement, convection, effusion combustor tile

Country Status (4)

Country Link
US (2) US10451276B2 (en)
EP (1) EP2965010B1 (en)
CA (1) CA2904200A1 (en)
WO (1) WO2014137428A1 (en)

Families Citing this family (21)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US9969654B2 (en) * 2014-01-24 2018-05-15 United Technologies Corporation Method of bonding a metallic component to a non-metallic component using a compliant material
EP3102884B1 (en) * 2014-02-03 2020-04-01 United Technologies Corporation Stepped heat shield for a turbine engine combustor
DE102014221225A1 (en) * 2014-10-20 2016-04-21 Siemens Aktiengesellschaft Heat shield element and method for its production
US10451280B2 (en) * 2015-02-16 2019-10-22 United Technologies Corporation Combustor panel having material transition region
US20160258623A1 (en) * 2015-03-05 2016-09-08 United Technologies Corporation Combustor and heat shield configurations for a gas turbine engine
US10480787B2 (en) 2015-03-26 2019-11-19 United Technologies Corporation Combustor wall cooling channel formed by additive manufacturing
US10208955B2 (en) * 2015-04-07 2019-02-19 United Technologies Corporation Ceramic and metal engine components with gradient transition from metal to ceramic
US10520193B2 (en) * 2015-10-28 2019-12-31 General Electric Company Cooling patch for hot gas path components
US10260750B2 (en) * 2015-12-29 2019-04-16 United Technologies Corporation Combustor panels having angled rail
GB201610122D0 (en) * 2016-06-10 2016-07-27 Rolls Royce Plc A combustion chamber
US10393380B2 (en) 2016-07-12 2019-08-27 Rolls-Royce North American Technologies Inc. Combustor cassette liner mounting assembly
US10386067B2 (en) * 2016-09-15 2019-08-20 United Technologies Corporation Wall panel assembly for a gas turbine engine
DE102017203326A1 (en) * 2017-03-01 2018-09-06 Rolls-Royce Deutschland Ltd & Co Kg Combustor shingle arrangement of a gas turbine
US10775044B2 (en) 2018-10-26 2020-09-15 Honeywell International Inc. Gas turbine engine dual-wall hot section structure
US11371702B2 (en) 2020-08-31 2022-06-28 General Electric Company Impingement panel for a turbomachine
US11460191B2 (en) 2020-08-31 2022-10-04 General Electric Company Cooling insert for a turbomachine
US11994292B2 (en) 2020-08-31 2024-05-28 General Electric Company Impingement cooling apparatus for turbomachine
US11994293B2 (en) 2020-08-31 2024-05-28 General Electric Company Impingement cooling apparatus support structure and method of manufacture
US11614233B2 (en) 2020-08-31 2023-03-28 General Electric Company Impingement panel support structure and method of manufacture
US11255545B1 (en) 2020-10-26 2022-02-22 General Electric Company Integrated combustion nozzle having a unified head end
US11767766B1 (en) 2022-07-29 2023-09-26 General Electric Company Turbomachine airfoil having impingement cooling passages

Family Cites Families (50)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
GB2049152B (en) * 1979-05-01 1983-05-18 Rolls Royce Perforate laminated material
US4302940A (en) * 1979-06-13 1981-12-01 General Motors Corporation Patterned porous laminated material
US4312186A (en) * 1979-10-17 1982-01-26 General Motors Corporation Shingled laminated porous material
US4296606A (en) * 1979-10-17 1981-10-27 General Motors Corporation Porous laminated material
US4628694A (en) * 1983-12-19 1986-12-16 General Electric Company Fabricated liner article and method
US4751962A (en) * 1986-02-10 1988-06-21 General Motors Corporation Temperature responsive laminated porous metal panel
DE3615226A1 (en) * 1986-05-06 1987-11-12 Mtu Muenchen Gmbh HOT GAS OVERHEATING PROTECTION DEVICE FOR GAS TURBINE ENGINES
US5195243A (en) * 1992-02-28 1993-03-23 General Motors Corporation Method of making a coated porous metal panel
US5363654A (en) * 1993-05-10 1994-11-15 General Electric Company Recuperative impingement cooling of jet engine components
US5528904A (en) * 1994-02-28 1996-06-25 Jones; Charles R. Coated hot gas duct liner
DE69526615T2 (en) * 1994-09-14 2002-11-28 Mitsubishi Jukogyo K.K., Tokio/Tokyo Wall structure for the outlet nozzle of a supersonic jet engine
DE19751299C2 (en) * 1997-11-19 1999-09-09 Siemens Ag Combustion chamber and method for steam cooling a combustion chamber
GB9803291D0 (en) * 1998-02-18 1998-04-08 Chapman H C Combustion apparatus
GB9926257D0 (en) * 1999-11-06 2000-01-12 Rolls Royce Plc Wall elements for gas turbine engine combustors
GB2373319B (en) 2001-03-12 2005-03-30 Rolls Royce Plc Combustion apparatus
EP1288578A1 (en) 2001-08-31 2003-03-05 Siemens Aktiengesellschaft Combustor layout
DE10155420A1 (en) 2001-11-12 2003-05-22 Rolls Royce Deutschland Heat shield arrangement with sealing element
US6701714B2 (en) * 2001-12-05 2004-03-09 United Technologies Corporation Gas turbine combustor
GB2384046B (en) * 2002-01-15 2005-07-06 Rolls Royce Plc A double wall combuster tile arrangement
US7093439B2 (en) 2002-05-16 2006-08-22 United Technologies Corporation Heat shield panels for use in a combustor for a gas turbine engine
FR2852003B1 (en) * 2003-03-04 2005-05-27 Snecma Propulsion Solide PROCESS FOR PRODUCING A MULTIPERFORATED PART IN CERAMIC MATRIX COMPOSITE MATERIAL
US7146815B2 (en) 2003-07-31 2006-12-12 United Technologies Corporation Combustor
US7363763B2 (en) 2003-10-23 2008-04-29 United Technologies Corporation Combustor
US7140185B2 (en) * 2004-07-12 2006-11-28 United Technologies Corporation Heatshielded article
US7464554B2 (en) 2004-09-09 2008-12-16 United Technologies Corporation Gas turbine combustor heat shield panel or exhaust panel including a cooling device
US7219498B2 (en) * 2004-09-10 2007-05-22 Honeywell International, Inc. Waffled impingement effusion method
GB0425794D0 (en) * 2004-11-24 2004-12-22 Rolls Royce Plc Acoustic damper
GB2420614B (en) 2004-11-30 2009-06-03 Alstom Technology Ltd Tile and exo-skeleton tile structure
EP1672281A1 (en) * 2004-12-16 2006-06-21 Siemens Aktiengesellschaft Thermal shield element
US7954325B2 (en) 2005-12-06 2011-06-07 United Technologies Corporation Gas turbine combustor
GB0601413D0 (en) * 2006-01-25 2006-03-08 Rolls Royce Plc Wall elements for gas turbine engine combustors
EP1813869A3 (en) 2006-01-25 2013-08-14 Rolls-Royce plc Wall elements for gas turbine engine combustors
GB0601418D0 (en) 2006-01-25 2006-03-08 Rolls Royce Plc Wall elements for gas turbine engine combustors
DE102006026969A1 (en) 2006-06-09 2007-12-13 Rolls-Royce Deutschland Ltd & Co Kg Gas turbine combustor wall for a lean-burn gas turbine combustor
DE102007018061A1 (en) 2007-04-17 2008-10-23 Rolls-Royce Deutschland Ltd & Co Kg Gas turbine combustion chamber wall
US7905094B2 (en) 2007-09-28 2011-03-15 Honeywell International Inc. Combustor systems with liners having improved cooling hole patterns
GB2453946B (en) 2007-10-23 2010-07-14 Rolls Royce Plc A Wall Element for use in Combustion Apparatus
FR2929690B1 (en) * 2008-04-03 2012-08-17 Snecma Propulsion Solide COMBUSTION CHAMBER SECTORIZED IN CMC FOR GAS TURBINE
US8661826B2 (en) 2008-07-17 2014-03-04 Rolls-Royce Plc Combustion apparatus
US20100095680A1 (en) 2008-10-22 2010-04-22 Honeywell International Inc. Dual wall structure for use in a combustor of a gas turbine engine
US9897320B2 (en) 2009-07-30 2018-02-20 Honeywell International Inc. Effusion cooled dual wall gas turbine combustors
GB0913580D0 (en) * 2009-08-05 2009-09-16 Rolls Royce Plc Combustor tile
US9416970B2 (en) 2009-11-30 2016-08-16 United Technologies Corporation Combustor heat panel arrangement having holes offset from seams of a radially opposing heat panel
US20110185739A1 (en) 2010-01-29 2011-08-04 Honeywell International Inc. Gas turbine combustors with dual walled liners
US8490399B2 (en) * 2011-02-15 2013-07-23 Siemens Energy, Inc. Thermally isolated wall assembly
US8978385B2 (en) * 2011-07-29 2015-03-17 United Technologies Corporation Distributed cooling for gas turbine engine combustor
US9057523B2 (en) * 2011-07-29 2015-06-16 United Technologies Corporation Microcircuit cooling for gas turbine engine combustor
US20140216042A1 (en) * 2012-09-28 2014-08-07 United Technologies Corporation Combustor component with cooling holes formed by additive manufacturing
EP2971971B1 (en) * 2013-03-13 2018-11-28 Rolls-Royce North American Technologies, Inc. Check valve for propulsive engine combustion chamber
US9879861B2 (en) * 2013-03-15 2018-01-30 Rolls-Royce Corporation Gas turbine engine with improved combustion liner

Also Published As

Publication number Publication date
US20200025378A1 (en) 2020-01-23
EP2965010B1 (en) 2018-10-17
US10451276B2 (en) 2019-10-22
US20140250894A1 (en) 2014-09-11
WO2014137428A1 (en) 2014-09-12
EP2965010A1 (en) 2016-01-13

Similar Documents

Publication Publication Date Title
US20200025378A1 (en) Dual-wall impingement, convection, effusion combustor tile
US10337737B2 (en) Combustor tile
US7306424B2 (en) Blade outer seal with micro axial flow cooling system
US7770397B2 (en) Combustor dome panel heat shield cooling
EP2549188B1 (en) Insert for a gas turbine engine combustor
CA2855776C (en) Interlocking combustor heat shield panels
US10648666B2 (en) Angled combustor liner cooling holes through transverse structure within a gas turbine engine combustor
US9982890B2 (en) Combustor dome heat shield
US10101029B2 (en) Combustor panels and configurations for a gas turbine engine
US10731855B2 (en) Combustor panel cooling arrangements
US11009230B2 (en) Undercut combustor panel rail
CA2920188C (en) Combustor dome heat shield
US20140000267A1 (en) Transition duct for a gas turbine
US10605169B2 (en) Combustor panel cooling arrangements
US11492911B2 (en) Turbine stator vane comprising an inner cooling wall produced by additive manufacturing
CA2854848C (en) Asymmetric combustor heat shield panels
US20140238027A1 (en) Thermally compliant dual wall liner for a gas turbine engine
US12098654B2 (en) Bi-cast trailing edge feed and purge hole cooling scheme

Legal Events

Date Code Title Description
EEER Examination request

Effective date: 20181115

FZDE Discontinued

Effective date: 20220603

FZDE Discontinued

Effective date: 20220603